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THE USE OF GRAPHITE/EPOXY COMPOSITE STRUCTURES IN SPACE APPLICATIONS
Rudy Lukez
Morton Thiokol, Inc, Aerospace Group
Brigham City Utah
ABSTRACT
Graphite/epoxy composite materials are being
used increasingly for numerous space applications.
Engineers are interested In these materials because
of their favorable mechanical characteristic of high
strength/high stiffness to weight ratio and potential
for zero or near-zero coefficient of thermal expansion.
This paper presents an overview of graphite/epoxy
composite use for space applications. The historical
uses and environmental concerns of graphite/
epoxy composites in the low earth orbit are
reviewed. Detailed information on the design, fabrication and testing of struts for potential Space Station use is presented as an example of graphite/
epoxy material usage for space applications.
GRAPHITEfEPOXY COMPOSITE STRUCTURES
WITH SPACE EXPERIENCE
Even though using graphite/epoxy structures for
space applications appears to be a novel idea,
there are a number of space vehicles which have
successfully employed composites. Published
papers contain numerous examples of these applications. A few spacecrafts which demonstrate the
use of graphite/epoxy composites are:
I. Intelsat IV: ThiS communications satellite was
built by Ford Aerospace and Communications
Corporation and is shown in Fig. I. It used a
hybrid horizontal cross arm made of graphite/
epoxy tape and aluminum. This design was
selected to meet requirement for small deflections and low weight. IRef. I)
INTRODUCTION
Vehicle and structure applications in space encounter a variety of design requirements which call for
new and creative material applications. Problems
associated with the space environment such as
radiation and atomic oxygen, physical demands
based on size and weight and performance goals
for longevity and funaionality present a variety of
challenges which call for new technologies.
Advances in materials during the past twenty
years have solved many of the challenges found in
the Space Program. One family of material systems, composites, has been used to meet many
varied space requirements.
One material system of particular interest is
graphite/epoxy composites. These materials have
been used to meet a wide variety of design parameters ranging from minimal weight to high structural stiffness requirements. Because of their ability
to meet many diverse design requirements simultaneously the use of graphite/epoxy composites is
expected to increase.
Figure 1. Intelsat Communications Satellite
(NASA photo)
2. Anik Anik, pictured In Fig 2, IS the Canadian
communications satellite and was bUilt by RCA
Astro-Electronics Anik's antenna structure is a
6 ft high elliptical parabola fabricated from
graphite/epoxy skin/aluminum honeycomb
core components. This material system was
chosen to satisfy Anik's requirements for:
1. controlling solar radiation effects,
2. reducing thermal distortions and
3. maintaining high structural stiffness.
(Ref. 1)
Figure 2. Anik, A Canadian Communications
Satellite (NASA photo)
3.
The Viking spacecrafL which was used
In the late 19705 for the exploration of Mars,
employed graphite/epoxy parts built by Ford
Motor Company's Philco Division. As with
Anlk, a 5 ft diameter antenna using graphite/
epoxy skins over an aluminum honeycomb
for the lander (see Fig. 3) and
Fig, 4). Composites were selected
to meet requirements for
1. low thermal response,
2. high structural stiffness and
3. low weight (Ref. I)
Figure 4. Viking Satellite Above Mars
Deploying Viking Lander (NASA photo)
Figure 3. Viking Lander Spacecraft on the
Surface of Mars (NASA photo)
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4. DSCS III: Another space vehicle which uses
graphite/epoxy composites is the third Defense
Satellite Communication System [DSCS III) satellite shown in Fig. 5. This geosynchronous communications satellite employs several graphite/
epoxy parts including: 1. the launch adaptor,
and 2. five different structural support strut configurations for primary and secondary applications. The use of graphite/epoxy parts reduced
the satellite's weight by 25 lb. [Ref. 2)
THE SPACE ENVIRONMENT AND COMPOSITES
The space environment IS not very conducive to
unprotected graphite/epoxy composite systems or
most other organic materials. Of particular interest
is the environment surrounding low-earth orbit
(LEO) which is where many orbital spacecrafts,
including the Space Shuttle and Space Station, are
designed to operate.
Specific environmental concerns include:
Radiation
Figure 5. Defense Satellite Communications
System III
There are many other examples from past current
and future space projects which call for graphite/
epoxy composite structures. Future planned composite use include:
1.
2.
3.
4.
5.
the Space Telescope depicted in Fig. 6,
cryogenic tanks,
advanced.antenna systems,
Space Tug,
Space Station, which is discussed in greater
detail below, and
6. Small, lightweight satellite systems.
Space radiation, comprised of electrons and protons in the earth's magnetic field, can alter the
dimensional stability and thermomechanical properties of graphite/epoxy composites; in fact any
part with an organic constituent can be adversely
affected by radiation if not properly protected.
Radiation produces damaging effects in organic
compounds, such as graphite/epoxy composites,
through chain-scission and cross-linking at various
radiation doses ranging from 105 to 109 rads. [Ref.
3) This causes a reduction of a composite's glass
transition temperature [T g).
Radiation damage effects include changes in structural
1.
2.
3.
4.
5.
6.
7.
melting and softening points,
hardness,
electrical properties,
ultimate tensile strength,
elongation,
modulus of elasticity and
dimensional instability. [Ref. 3)
Since radiation effects are cumulative, structures
planned for long service life need adequate protection from radiation exposure. The Space Station will
see the equivalent of 15 years of radiation exposure during its planned life.
Thermal cycling
Most spacecraft missions involve orbits around the
earth. During these orbits, a spacecraft passes in
and out of the earth's shadow and experiences
extreme minimum and maximum temperatures on
a cyclical schedule. This cycling can induce microcracks in composite resins which reduce part stiffness and changes the overall coefficient of thermal
expansion (CTE). (Ref. 4)
Thermal cycling effects can be prevented by using
protective coatings over exposed composite parts.
Typical protective coatings include aluminum foils.
The purpose of these protective coatings is to distribute temperatures uniformly around and along
the exposed structure.
Figure 6.
photo)
The Space Telescope (NASA
As structures are scheduled for increased service
life in space, thermal cycling will become more
important. As an example of the magnitude of this
cycling, the
Station during its 30 year life will
experience 175,000 cycles. (Ref. 5)
Atomic oxygen
AtomiC oxygen IS the major constituent in the LEO
environment between 200 km and 500 km. It is
also now considered a significant concern when
spacecraft are deSigned for near earth applications.
(Ref. 5)
The effects of atomic oxygen were observed on
early Space Shuttle flights. Experiments on
Shuttle (STS) flights 5, 8, and 4J-G showed that
non-metals, including graphite/epoxy composites,
are very reactive to atomic oxygen while metals
are nearly unaffected, Graphite/epoxy systems
react with atomic oxygen to form volatile oxides
which cause the matenal to dissolve. Given sufficient time, atomic oxygen could completely dissolve
a composite part in LEO,
Atomic oxygen degradation is so severe that one
paper noted:
"For many years, the assumption was that
the aspects most degrading to materials in
the low earth orbital environment were ultra-
Violet radiation and thermal vacuum exposure. From the results of early experiments
conducted on the
Shuttle, it now
appears that atomic oxygen effects will by far
be more damaging." (Ref. 5)
To prevent damage to graphite/epoxy parts in
space, LEO exposed composites need protective
coatings which are non-reactive with atomic oxygen The Space Station, for example, will require
aluminum foil, probably 4 to 5 mils, over the
graphite/epoxy struts used in the connecting truss.
This concept IS simple and economical; in addition,
besides atomic oxygen, aluminum foil should provide long term protection for other environmental
problems.
GRAPHITE/EPOXY STRUTS FOR SPACE STATION
APPLICATIONS
Earlier Space Station requirements based on the
proposed dual keel rectangUlar truss structure,
shown in Fig. 7, called for nearly 23,000 ft of
graphite/epoxy struts. Updated baseline concepts,
such as the version shown in Fig. 8, will require
fewer struts initially, but could still be easily
expanded to the earlier proposed structure.
Figure 7, Space Station Dual Keel 5-meter
Truss Configuration (NASA photo)
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Figure 8.
Current Space Station Phase I Baseline Design. The Horizontal Truss Will Be
Made From Graphite/Epoxy Struts to Which Laboratory and Habitation Modules and Solar
Panels are Attached (NASA photo)
Selecteej fa:
stiffness, minimum thermal distortion and low weight characterIStics, graphltel
epoxy components wi,: be required ro Survive 30
years in low earL'l orbit 250 nautical miles above
t:'le earth's surface. Construction of the
Station is scheduled for ,he early to mid 1990s. 500
years after ChrIStopher Coiumbus sailed to tile
New World.
To meet the requirements for the Space Station
truss structure. a program to design, fabricate and
test graphite/epoxy composite struts was implemented. Components of this program are
explained below
Requirements
Major Space Station design requirements set forth
in early 1985 were: fRef. 6)
• Dimensional stability: Dimensional stability is
one of two key design drivers that dictate the
use of low coefficient of thermal expansion
fCTE) matenals for the struts. Because of their
low weight. graphite/epoxy composites are
favored over metals for the truss struts. Satisfying
the dimensional stability reqUirement ensures
trouble-free assembly of the truss and maintains
pointing and tracking accuracy of the on-board
experiments and power generation equipment
during thermal exposure.
• Axial stiffness: Axial stiffness is the other key
design driver affecting the selection of materials
for the Space Station truss structure. An analysis
of the overall structure's flexural
and ta:stiffness reqUirements leadS to a
slonal
design
by longitudinal modulus IE 1)
of the struts.
• Strength: Strength :s not a
design factor
for the Space Station struts due to the expected
low operational loads. The area of most interest
is the transition region between the composite
strut and end fittings where the highest loads
are anticipated.
• Column stability: Column stability also is a concern because of the effect end fittings have on
end fixity rather than the magnitude of the loads
involved.
• Age life: Age life in space is a major materials
selection factor and creates concerns with the
use of composites because of atomic oxygen
and thermal cycling.
• Atomic oxygen: Since atomic oxygen particles
degrade the epoxy in a graphite/epoxy system
and thermal cycling causes microcracks to form
in the epoxy due to CTE mismatch between the
graphite fiber and epoxy resin, protective metallic coatings and "toughened" resins are necessary.
• Damage resistance and repair: Damage resistance and repair are practical requirements necessary for the low cost implementation of composites to the Space Station. Damage resistance
is a function of material selection, fiber orientation and external protection.
Oeslgn and analysIs
Using laminated plate theory !LPT), numerous fiber
orientations and material systems were evaluated;
there are a number of materials and fiber orientations which satisfy the proposed requirement with
weight and cost the obvious trade-off, Sample
designs are listed in Table I.
Since mechanical requirements for Space Station
struts are rather straight forward, elementary hand
c;alculations and LPT evaluations provided sufficIent design and analysis information.
As composite studies continued, the issue of
potential thermal cycling induced matrix microcracking required a design change. Resin microcracking occurs readily in composites with high
angle cross-ply lay-ups, such as a [± 7s o ml
± ISoPnls combination, when exposed to the
temperature extremes typical of low earth orbit.
Composites with lower angle cross-ply lay-ups are
less susceptible to microcracking due to lower
interlaminarthermal stresses. So, [± 4s o/0 04J2 and
[±4So/± Iso/0021s lay-ups using .005 in. thick
plies were selected for further studies
Table 1
SAMPLE LONGITUDINAL CTE AND MODULUS VALUES
FOR P75/934 GRAPHITE/EPOXY TUBE
DESIGNS
Layup
[± 75/± 155]S
[ ± 751 ± 152]S
[± 101 ± 305]S
[± 101 ± 302]S
[± 751 ± 452]S
[ ± 451 ± 1 55]S
[± 451 ± 1521S
[± 45/041S
[ ± 451 ± 15/02]S
Ply
Count
Ply Thickness
(in. )
CTE x 10~6
(in.lin. OF)
24
12
24
12
24
24
12
12
12
0.0025
0.005
0.0025
0.005
0.0025
0.0025
0.005
0.005
0.005
~0.53
-0.37
~1.22
~1.22
-0.51
-0.99
-0.88
~0.72
-0.96
Modulus
(msi)
28.4
23.2
26.6
26.6
25.4
23.7
25.8
27.6
24.6
BB253·1C
Strut fabrication studies
A variety of fabrication methods were studied during the initial development effort, A matrix comparing the advantages and disadvantages of several
fabrication methods is shown in Table 2.
• Transverse tape rolling: The selected process
was transverse tape rolling because of its layup
versatility. low void content and, if automated,
very low cost. Ply orientations from 0° to 90° in
numerous combinations and consistent void
contents below .5 percent have been demonstrated using this process. This process, shown
in Fig. 9, was chosen because of its outside
dimensional control capability, ease of use and
high quality surface finish. The machine used in
this process is pictured in Fig. 10.
• Filament winding: Filament winding was not
selected because it is not well suited for long
struts with small diameters. Morton Thiokol
studies show it is inefficient to filament Wind
parts with a length/diameter ratio greater than
J 0 when low angle orientations lIess than
± 30°) are used due to equipment constraints.
The low wind angles would be required to meet
minimum CTE requirements for the struts.
• Braiding: Because of numerous fiber crossovers
and the program's low void content requirements, braiding was not selected. Typically,
braided parts have higher void contents and
lower fiber volumes than unidirectional tape.
Braided parts can range from 0 to 8 percent
voids while unidirectional tape parts vary from 0
to 2 percent voids. Also, when the original fabrication trade studies were completed, high modulus fibers (above 40 msil could not be braided
without considerable damage.
• Pultrusion: Pultrusion processing could net high
production rates at 5 to 15ft per minute but was
not selected because of high equipment costs,
problems with using epoxy resins and lack of ply
angle versatility.
• Convolute strut winding: As with pultrusion
methods, convolute strut winding could reach
high production rates but was not selected
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Table 2
COMPARISON OF SPACE STATION STRUT FABRICATION METHODS
Tape Rolling
•
Filament
Winding
•
•
•
•
•
• • • •
• • • •
•
Pultrusion
Convolute
Winding
•
•
•
•
•
• • • •
•
•
•
•
•
•
•
•
•
•
•
• •
Braiding
•
• •
•
•
88218-lG
•
Favorable
Capability
Notes: 1. Ability to use different material systems
2. Ability to easily apply pr<?tective overwrap materials as aluminum foil
1. Cut Tape Sections
2. Ply Tape Layers Together
[Q]1Ql[Q]1Q]
[Q][QI[QJIQ]
[Q][QJIQl[QJ
4. Insert Into Mold
1Ql[Q][QJ[QJ
[QIIQ]IQ][QJ
IQ]IQ][QJ[Q]
5. Install Into Batch
Cure Stand
6. Batch Cure
'f
7. Remove From Mold
Figure 9.
8. Cut Ends to Length
9. Inspect Finished tube
Space Station Strut Fabrication Process
Figure 10. Strut Rolling Machine Designed and Built By Morton Thiokol, Inc.
because of expensive equipment requirements.
In addition, the process IS not well suited for
flexible fiber OrIentation and layup variations
which IS important to develop the optimum strut
layup.
Strut fabrication process
To fabricate a
Station, strut sheets of preimpregnated tape were cut and laid to a specific ori-
entation as determined by the design; a sample
layup is shown in Fig. liThe layup was then
placed on the feed table of the strut rolling
machine and fed Into the machine to be rolled up
on the mandrel. The tape was then automatically
pulled Into the machine and rolled under controlled compaction. This process was very quick
and required only a few mandrel revolutions to roll
the strut
I
120 0
Sheet 3
,
0
120 0
f---
(a). Feed Sequence Onto Male Mandrel
+A
~A
±B
±B
+B
±B
±B
-line of Symmetry- - - -
1
J Sheet 1
j
Sheet 2
B
00
2400
#3
#2
1200
(ci. Sheet Orientation on
Completed Tube
: : } Sheet 3
(bi. Sheet Construction
Figure 11.
Sample Graphite/Epoxy Tape Layup for Space Station Struts
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The rolled strut was then placed in a female mold,
stacked in a rack, and placed in an oven for curing.
After cure and cool down, the strut was pulled
from the mold. This was easily done by hand since
the CTE difference between the mold and strut
allowed for"suffiClent clearance.
completed or are in progress to verify the ability of
graphite/epoxy struts to meet Space Station requirements. The programs are:
• Mechanical testing: Room temperature tension and compression tests were completed
on 26 specimens. The maximum load used
was approximately 2,500 Ib, which is typical
of operational loads in an orbiting space
structure and well below the strength capability of the struts. LPT design analyses determined the theoretical modulus of elasticity for
each lay-up tested.
Once the strut was removed, the ends were
parted to length. Tests showed tha~ void content
was less than .5 percent with a finished straightness jper ANSI Y 14 5J of less than .03 percent per
10ft length. Additional tests measured fiber volumes of 60 percent surface finishes of 8 rms, and
outside diameter tolerance of ± .001 in.
• Table 3 shows results obtained from ten
[±45°/0041s and [±45°/± 15°10 0 21s layups. The data, obtained from specimens
made from one lot of material, slightly
exceeded theoretical predictions.
With co-cured protective coatings, it was possible
to apply various foils to the inside and outside of a
strut without additional processing. To apply an
aluminum foil coating to the outside, an aluminum
sheet was rolled up with the strut as the last wrap
during the rolling process. It was then placed in the
mold and cured. The mandrel during cure forms
the foil and strut tightly together to produce a voidless bond.
• Short-term thermal cycling: In 1988, tension,
compression and CTE tests are planned for
short-term thermal cycled specimens. The
thermal proposed tests will simulate the LEO
environment 1,000 times.
Strut testing
• Long-term thermal cycling: Long-term thermal cycling tests will begin soon at the Center
Several different test programs have either been
Table 3
COMPARISON OF TWO GRAPHITE/EPOXY
STRUT DESIGN MECHANICAL TESTS
Test
Number
P75S/934
[±45°/± 15 %0 2]S
1
2
3
4
5
6
7
8
9
10
Average
P75S/934
[±45 % 04]S
1
2
3
4
5
6
7
8
9
10
Average
Tension
Modulus
(msi)
Compression
Modulus
(msi)
24.7
25.5
26.0
25.2
27.7
26.6
26.0
24.7
26.6
26.9
26.0
23.8
24.6
25.1
24.2
26.7
25.6
25.1
23.9
25.6
25.7
25.0
31.8
29.9
25.9
28.1
28.0
28.8
30.1
30.6
33.0
29.8
29.6
30.7
28.9
25.0
27.1
27.0
27.9
29.1
29.6
31.7
29.2
28.6
Predicted
Modulus (msi)
24.6
27.6
88253-10
for Space Engineering. Utah State University,
Logan, Utah under a JOint program with
Morton Thiokol, Inc. The test will subject strut
samples to a -J50o/2JooF environment
10,000 times uSing the fixture shown in Fig.
r 2.
tower for large space structure simulations. This
deployab'e tower will be transported to and
frnm nrhit viA thp <;n.::lrp <;hllttlp
• End fittings: Different concepts for integrating
end fittings with struts during fabrication are
being designed, fabricated and tested. Integrated end fittings would offer reduced fabncation cost and increased joint reliability,
Heating
Coil
Aluminum
Receptacle
(Top and Bottom)
.'
Nitrogen
Gas
Filled
Low
Conductivity
Tube
L(Approx)
200
'
.
..
Test
Specimen
Liquid
Nitrogen
Channel
(-3200FI
Figure 12. Long Term Thermal Cycling Test
Fixture Designed and Built at Utah State University
Figure 13. Deployable Mast Subsystem
(Photo Astro Aerospace Corporation)
SUMMARY
The use of graphite/epoxy composite parts for
space applications is already well established.
USing graphite/epoxy parts for space vehicles
and structures has many advantages including:
I . critical weight savings,
2. improved control of thermal distortions,
3. increased structural stiffness.
The strut process developed for Space Station
applications readily lends itself to other projects.
Diameters. lengths. and layups can easily be
varied depending on the design.
Spacecraft designers also must consider the
severe space environment when specifying
graphite/epoxy materials for their vehicles and
structures. Radiation, atomic oxygen, and thermal cycling effects must be considered. Fabrication techniques do exist however, which can
easily protect graphite/epoxy material systems in
the space environment.
For example. the design, fabrication, and test
experience described above has been used in
developing preliminary concepts for the Deployable Mast Subsystem (OMS). The OMS, shown
in Fig. J 3 and designed by Astro Aerospace
Corporation and Harris Corporation, uses small
diameter (.5, .75. and J.O in.) graphite/epoxy
struts of different lengths to build a 180 ft long
Rolled composite struts designed for the proposed Space Station demonstrate that graphite/
epoxy material systems readily can be used to
meet the program's various mechanical and
environmental requirements. At the same time,
the Space Station struts and the associated fabrication process could be used for other spacecraft applications.
Other Strut Uses
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ACKNOWLEDGEMENTS
The author thanks Volker Teller, Ed Mason, Pete
Evanoff and David R. Nelson of Morton Thiokol,
Inc., and Harley J. Rockoff of Rockwell International's Space Station Systems Division for their
assistance in preparing this paper.
ABBREVIATIONS
CTE ........ Coefficient of thermal expansion
OMS ........... Deployable mast subsystem
LEO ..................... Low earth orbit
LPT ................ Laminated plate theory
REFERENCES
I. Mayer, N.J., Carbon Composites in Space
Vehicle Structures, International Conference
on Carbon Fibers, Paper No. 39, London,
1974.
2. Mazzio, V F. and C. H. Bixler, Optimized
Design and Fabrication Process for
Advanced Composite Spacecraft Structures,
American Institute of Aeronautics and Astronautics, Paper No. A79-19616, New York,
1979.
3. Mauri, R. E. and Crossman, F. W .. Space
Radiation Effects on Structural Composites,
American Institute of Aeronautics and Astronautics, Paper No. A83-16808, New York,
1983.
4. Tompkins, Stephen S, Sykes, George F., and
Bowles, David E, The Thermal and Mechanical Stability of Composite Materials for Space
Structures, IEEE/ASM/ASME/SME Space Technical Conference, Anaheim, California,
1985.
5. Leger, L., Visentine, J., Santos-Mason, B.,
Selected Materials Issues Associated With
Space Station, SAMPE Quarterly, Vol. 18, No.
2, January 1987, pp 48-54.
6. Lukez, R., Nelson, D., Teller, V, and Rockoff,
H., Design, Fabrication, and Testing of Rolled
Carbon/Epoxy Struts for Space Station Applications, 19th International SAMPE Technical
Conference, Washington, D.C., 1987
BIOGRAPHY
Rudy Lukez is a design engineer in Morton
Thiokol's Advanced Composite Structures Section. He is providing technical direction for a
variety of composite strut development programs currently being pursued at Morton
Thiokol's Aerospace Group. Mr. Lukez earned
his B.S. in Mechanical Engineering at Cleveland
State University, Cleveland, Ohio, in 1983 and
has worked for Morton Thiokol since 1984.
Prior to his employment at Morton Thiokol, Mr.
Lukez worked for Hughes Aircraft Company's
Radar Systems Group in EI Segundo, California.